Interior noise control is a primary concern in the operation of helicopters. While there are numerous sources of noise-generating vibrations occurring in an operating helicopter, such as the main rotor assembly, the main transmission, the engines, the tail rotor assembly, the hydraulic system, aerodynamic forces, etc., the vibrations emanating from the main rotor assembly and the main transmission have the most pronounced effect on helicopter interior noise levels, i.e., in the cockpit and/or cabin.
The operation of the main rotor assembly during forward flight of the helicopter results in the generation of "low frequency" vibrations at the blade passing frequency (and harmonics thereof). The blade passing frequency is equal to the product of the number of main rotor blades and the rotational speed of the rotor. For example, the main rotor assembly of the BLACK HAWK.RTM. helicopter (BLACK HAWK is a registered trademark of Sikorsky Aircraft), which has four main rotor blades, has a blade passing frequency of approximately 17 Hz. Similarly, the S-76.RTM. helicopter has a blade passing frequency of approximately 19 Hz.
The operation of the main transmission of the helicopter, in contrast, results in the generation of "high frequency" vibrations. For example, Sikorsky helicopters of the S-76.RTM. series, e.g., S-76A, S-76B, S-76C, have a main transmission that includes three stages of reduction gearing: a first stage for each engine output consisting of helical gearing, an intermediate stage consisting of spiral bevel gearing, and a final reduction stage comprising a central bull gear that intermeshes with right and left hand bull pinions (to combine the inputs of the two engines that provide the motive power for S-76.RTM. helicopters). Research has shown that the cockpit and/or cabin noise levels of S-76.RTM. helicopters are primarily the result of vibrations originating in the main transmission.
Narrow band Fast Fourier Transform analyses, A-weighted octave levels, and overall dBA levels recorded in the cockpits and/or cabins of S-76A, S-76B, and S-76C helicopters indicate that interior noise levels are predominantly the result of vibrations occurring at the bull gearing meshing frequency of 778 Hz, as illustrated in FIG. 1. The vibrations produced by the first and second reduction stages of S-76.RTM. main transmission gearboxes, i.e., noise levels generated by the helical and spiral gearing as illustrated in FIG. 1, occur at higher frequencies and typically are not significant relative to the dominant noise levels produced by the fundamental and first few harmonics of the bull gearing meshing vibrations.
The high and low frequency vibrations emanating from the main transmission and the main rotor assembly, respectively, are coupled into the helicopter airframe via a main transmission gearbox support truss or the main transmission gearbox (the feet thereof being mechanically secured to structural members comprising the airframe). The high and low frequency vibratory forces coupled into the airframe structure induce vibratory responses of many airframe natural modes. The natural modes of the airframe can excite natural modes of the cockpit and/or cabin acoustic volume. It is the response of the acoustic modes that produces the undesirable noise levels within the helicopter cockpit and/or cabin.
Such noise levels generally cannot be effectively abated by parasitic type acoustic treatment of the cockpit and/or cabin interior. Acoustic panels or blankets may be partially effective for very high frequency induced noise, but are not very effective vis-a-vis induced noise in the 300 to 1000 HZ range and don't work at all with respect to low frequency induced noise emanating from the main rotor assembly. The weight penalty incurred by the use of acoustic panels or blankets, moreover, negatively impacts the performance capability of the helicopter.
Another passive technique involves the use of vibration isolators at the interface between the main rotor assembly/transmission gearbox and the airframe structure. Such vibration isolators transmit only a reduced portion of the noise-generating vibrations into the helicopter airframe due to their inherent softness. These vibration isolators, however, must be interposed in the primary load path of the helicopter, and transmission deflections under steady flight loads may cause high speed engine-to-transmission drive shaft deflections that may adversely impact shaft reliability.
Active noise suppression methods are known for alleviating the noise generated by high and/or low frequency vibrations emanating from the main transmission and/or the main rotor assembly. U.S. Pat. No. 4,819,182 discloses an apparatus and method primarily for reducing induced vibrations in a helicopter fuselage due to low frequency vibrations emanating from the main rotor assembly. The embodiment, described in the '182 application, includes a raft structure, which supports the main transmission gearbox, that is attached to the helicopter fuselage by elastomeric units that provide the load path for coupling main rotor assembly loads into the fuselage. Electro-hydraulic actuators are disposed between the raft and the fuselage adjacent respective elastomeric units and are operative in response to control signals to produce reactive forces that reduce vibrational loads induced in the helicopter fuselage by the main rotor assembly. Accelerometers mounted in the helicopter cockpit and cabin measure the vibratory loads and provide control signals to control the magnitude and phase of the reactive forces provided by the actuators.
U.S. Pat. No. 4,600,863 describes an actuator for reducing the coupling of vibrations from a source to the source support structure. The actuator is operative in response to signals provided by an accelerometer to provide vibration-canceling forces along the axis of the actuator. An in-line shear isolator is provided in series with the axis of the actuator and is operative to decouple source-derived vibrations that are orthogonal to the axis of the actuator. The in-line shear isolator allows a single actuator to be utilized to reduce the effect of source-derived vibrations on the support structure.
U.S. Pat. No. 4,795,123 describes an active damping system wherein a plurality of damping elements are operative to dampen sensed vibrations in a mechanical structure over the range of 20 to 15,000 Hz at respective structure contact points. Each damping element, which includes an accelerometer, an integrator, a mixer, a power amplifier and a driver, is operative to dampen vibrations in one dimension along the axis formed by the element and its structural contact point. Damping of vibrations in three dimensions at a specified point is accomplished by attaching three damping elements at right angles to one another. One described embodiment includes an accelerometer coaxially disposed with respect to a wrap around driver while another described embodiment includes an accelerometer physically separated from the driver. One object of the vibration damping system described in the '123 patent is to dampen vibrations in mechanical structures while minimizing the impact on the design and operation thereof.
Active noise control systems that are designed to counteract low frequency vibrations emanating from the main rotor assembly, e.g., the apparatus described in the '182 patent, are generally not effective to counteract the high frequency vibrations emanating from the main transmission gearbox. And, while active damping systems such as that described in the '123 patent which are active over a wide frequency range that encompasses both the high and low frequency vibrations being coupled into the helicopter airframe may be utilized to attempt to counteract such vibrations, such a broadband system would generally not be effective in significantly reducing the high frequency vibrations emanating from the main transmission gearbox. The damping elements, being designed for broadband operation, are not functionally optimized for operation at one or more of the predetermined frequencies defining the high frequency vibrations emanating from the main transmission gearbox. The number of active damping systems of the type described in the '123 patent may be excessive since the number of systems must be equal to the number of natural vibrational modes of the helicopter airframe that contribute to excessive interior noise levels.
A need exists for an active control system that is operative to effectively cancel or nullify the high frequency vibrations emanating from the main transmission gearbox. The active noise control system should be design optimized for operation with respect to one or more of the identified high frequency vibrations emanating from the main transmission gearbox. The active noise control system should be design optimized to minimize the number of actuators required to effectively nullify the one or more identified high frequency vibrations emanating from the main transmission gearbox. The active noise control system should be further design optimized to eliminate induced contaminating vibrations arising from the operative interaction between the actuators and sensors of the system.